Typically, gas turbine engines include a compressor for compressing air, a combustor for mixing the compressed air with fuel and igniting the mixture, and a turbine blade assembly for producing power. Combustors often operate at high temperatures that may exceed 2,500 degrees Fahrenheit. Typical turbine combustor configurations expose turbine blade assemblies to these high temperatures. As a result, turbine blades must be made of materials capable of withstanding such high temperatures. In addition, turbine blades often contain cooling systems for prolonging the life of the blades and reducing the likelihood of failure as a result of excessive temperatures.
Typically, turbine blades are formed from a root portion having a platform at one end and an elongated portion forming a blade that extends outwardly from the platform coupled to the root portion. The blade is ordinarily composed of a tip opposite the root section, a leading edge, and a trailing edge. The inner aspects of most turbine blades typically contain an intricate maze of cooling channels forming a cooling system. The cooling channels in a blade receive air from the compressor of the turbine engine and pass the air through the blade. The cooling channels often include multiple flow paths that are designed to maintain all aspects of the turbine blade at a relatively uniform temperature. Some blades utilize a 5-pass serpentine arrangement in which cooling flows are routed span wise and distributed to forward, mid-chord and trailing edge sections of the blade.
The advancement of thermal barrier coating (TBC) technology has impacted the traditional cooling systems for turbine blades. As more industrial turbine blades are applied with thicker and lower conductivity TBC, the cooling flow demand for the TBC covered blade is reduced. As result, there is insufficient cooling to split the total cooling flow into two or three flow circuits and utilize a forward flowing serpentine cooling system. Cooling flow for the leading edge and trailing edge has to be combined with a mid-chord flow circuit to form a single 5-pass flow circuit. However, for a forward 5-pass flow circuit with total blade cooling flow, back flow margin (BFM) can become a design issue.
In a typical 5-pass aft flowing serpentine cooling design, the leading edge if cooled with backside impingement cooling together with leading edge showerhead film cooling. Cooling air is fed by the first up pass of the 5-pass serpentine flow channel. The main body is cooling by the serpentine flow channel with built in trip strips on the internal walls for the augmentation of internal heat transfer performance. The trailing edge is typically cooled with a double impingement cooling system in conjunction with pressure side bleed cooling. The blade tip is often cooled by bleed off from the serpentine turns at the tip end.
However, with the lower cooling flows utilized in more advanced TBC covered blades, the ability to use the impingement cooling mechanism with a pressure side bleed is compromised. Thus, there is a need for an effective cooling system for a low cooling flow environment in a hollow airfoil, such as a TBC enhanced blade.